RM A50K07
NACA
RESEARCH MEMORANDUM
AERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING SWEPT BACK 63° - EFFECTIVENESS AT SUPERSONIC SPEEDS OF A 30-PERCENT CHORD, 50- PERCENT SEMISPAN ELEVON AS A LATERAL CONTROL DEVICE
By Robert N. Olson and Merrill H. Mead
Ames Aeronautical Laboratory Moffett Field, Calif.
NATIONAL ADVISORY COMMITTEE FOR AERON AU TICS WASH INGTON
January 18, 1951 Declassified April 8, 1957
1 NACA RM A50K07
NATIONAL ADVISORY COMMITTEE FOR AERONATITI CS
RESEARCH MEMORANDUM
AERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING
0 SWEP"r BACK 63 - EFFECTIVENESS AT SUPERSONIC SPEEDS OF A
30-PERCENT CHORD, 50-PERCENT SEMISPAN ElEVON
AS A LATERAL CONTROL DEVICE
By Robert N. Olson and Merrill H. Mead
SUMMARY
The effectiveness of a 5O-percent-semispan, constant-percent-chord elevon and of uppe~urface spoilers as lateral-control devices for a wing-fuselage combination employing a wing swept back 630 has been determined experimentally for the range of Mac.h numbers from 1.2 to 1.7 at a Reynolds number of 1.5 million. The results are compared with calculated values of the rolling-moment-effectiveness parameter CZ Q as obtained by the application of the linearized theory of supersonlc flow.
For the elevon, results indicate that only about half of the rolling-moment effectiveness predicted by linear theory was realized. The variation of Mach number had little effect on the rolling-moment effectiveness of the elevon over the range investigated (Mach numbers of 1.2 to 1.7). Increasing angle of attack in the range of 00 to 100 had a moderate, adverse effect on the rolling-moment effectiveness of the elevon at a Mach number of 1.2; however, this effect diminished wi th increasing Mach number and became inappreciable in the range of 1.5 to 1.7 Mac.h number. The lateral-control characteristics of the elevon were little affected by variation of sideslip angle in the range from 50 to _5°. Elevon deflection produced small adverse yawing moments at positive angles of attack, w.hich were of suc.h small magnitude as to be of little concern.
The upper-surface spOilers, having a maximum projection equal to the maximum wing-section thickness, were found to be inferior to the elevons for lateral control of this model because of a rapid loss in 0 effectiveness above an angle of attack of 4 •
J 2 NACA RM A50K07
The present elevon was found to be more than adequate for rolling control of this wing-:fuselage combination if the wing were rigid; hovi ever, aeroelastic effects, computed from an approximate strip theory, were found to influence the rolling effectiveness of the elevon to the extent that structural considerations would be of prime importance in the final estimation of required elevon size.
INTRODUCTION
An investigation was made to determine means of attaining adequate lateral and longitudinal control for a wing-fuselage combination designed for efficient flight at supersonic speeds. This wing-fuselage combination, designed according to the theory advanced by J ones in ref erence 1, employed a wing of aspect ratio 3.5, taper ratio 0.25, and 630 sweepback of the leading edge. The results of wind-t unnel tests a t supersonic speeds to determine the effectiveness of a 3O-percent chord, 5O-percent-semispan elevon as a longitudinal control for this wing fuselage combination are presented (along with drag and hinge-moment data) in reference 2. The present report presents the results of a wind-tunnel investigation at supersonic speeds to determine the lateral control characteristics of this elevon. The lateral-control character istics of upper-surface, 5O-percent-semispan, outboard spoilers were also investigated.
NOTATION
All moment coefficients were referred to the stab ility axes wi th the origin located at the quarter-chord point of t he mean aerodynamic chord projected to the fuselage center line.
rolling-moment coefficient (rolling moment) qSb
damping-moment coefficient in roll; the rate of change of rolling moment coefficient C1 with wing-tip helix angle pb/2V, per radian
C elevon rolling-moment-effectiveness parameter for constant angle of 1 o attack(dC 1 ) do a. Cn yawing-moment coefficient referred to quarter-chord point of mean aerodynamic chord (yaWing moment) qSb NACA RM A50K07 3
M Mach number (Via)
R Reynolds number (pvc/~)
S wing area~ square feet
V airspeed, feet per second a speed of sound, feet per second b wing span~ feet b/2 (1 2 c dY), c wing mean aerodynamic chord 0 :feet b/2 f c dy o c local wing chord measured parallel to plane of symmetry, feet p angular velocity in roll, radians per second pb wing-tip helix angle, radians 2V q dynamic pressure (~V2), pounds per square :foot
~ angle of attack o:f :fuselage center line~ degrees
~ angle o:f sideslip, positive with right wing leading~ degrees
5 angle between wing chord and elevon chord, measured in a plane perpendicular to the elevon hinge line~ positive :for downward deflection with respect to the wing, degrees p mass density o:f air, slugs per cubic :foot
~ viscosity o:f air, slugs per foot-second
Coefficients indicated with a prime (.) are uncorrected :for tunnel pressure gradient and flow inclination.
APPARATUS AND TEST METHODS
The investigation was conduct~d in the Ames 6- by 6-foot supersonic wind tunnel which is equipped with an asymmetric adjustable nozzle per mitting a variation in Mach number from 1.2 to 2.0 in the pressure range 4 NACA RM A50K07 of 2 to 20 pounds per s'luare inch absolute. However, due to certain vibration di:fficulties of the present model support system, the Mach number range is temporarily limited to a maximum of 1.7.
The strain-gage balance and instrumentation are described in detail in reference 2.
The model, pertinent dimensions of which are presented with a plan form sketch in figure 1, is shown mounted in the test section in the photograph of figure 2. The wing was untwisted, had 00 incidence, and was composed of NACA 0010 airfoil sections perpendicular to the leading edge. A ~ercent-chord elevon was mounted on the outboard 50 percent of the right wing panel. Spoilers, in heights up to the maximum wing section thickness, were constructed of aluminum alloy angles and were located successively at 40-, 50-, and 6o-percent-chord stations of the outboard half of the right wing panel. The wing and elevon were of solid-steel construction, while the fuselage was of hollo~teel con struction to permit the installation of the four-component strain-gage balance. A cutaway schematic drawing showing this installation is pre sented in reference 2. The fuselage was designed on the basis of mini mum wave drag for a given volume and length and had a fineness ratio of 12.5, including that portion of the body cut off to provide a method of attaching the model to the sting support.
To shorte~ the testing time necessary for the lateral-control effectiveness investigation, the model was mounted in the test section with the plane of the wing placed vertically. With the model mounted in this manner, the present model support system permits continuous variation of the angle of sideslip during the test. Changing angle of attack, however, entailed changing the model support sting for each desired angle of attack.
Side forces, rolling moments, and yawing moments were measured over a Mach number range from 1.2 to 1.7 at a Reynolds number of 1.5 million. The data were then transferred to stability axes and reduced to final coefficient form. Force measurements were made at nominal angles of sideslip of _50, -2.5°, 00 , 2.50 , and 50 for angles of attack of 00 , 0 0 5 1 and 10°. The elevon deflection was varied in 10 increments from 10° to -300 •
The angle of sideslip of the model was determined optically by means of a cathetometer. The control-surface deflections, however, could not be determined during tests but were measured under stat ic conditions before each test and corrected for deflection under aero dynamic loading, using constants determined from static loading tests .
The angle of attack was fixed for each test condition by us ing a sting bent to the desired angle. Since the angles of attack could not ------
NACA RM A50K07 5
be measured during the test, the angle of attack was measured initially under static conditions and corrected for deflection under aerodynamic load, using constants determined from static loading tests.
REDUCTION OF DATA
Since the force and moment strain gages were located inside the model, there were no direct foree tares due to aerodynamic forces on the sting. Effects of support interference, for the present investi gation, are confined to changes in base pressure and thus an effect on drag. (See reference 3.)
Although the variations in stream curvature existing in the test section of the Ames 6- by 6-foot wind tunnel (discussed in reference 4) are such that lateral characteristics of a highly swept wing-body cam bination such as the present configuration, tested at Mach numbers other than 1.4, would be subject to some doubt, values of the incremental aerodynamic coefficients due to elevon deflection are reliable. Be cause of the axial variations in stream angle at Mach numbers above and below 1.4, however, it is difficult to determine the effective model angle of sideslip when testing with the plane of the wing vertical. Therefore, the angles of sideslip referred to in the report are nominal angles referenced to the borizontal plane and as such restrict the significance of the angle-of-sideslip eff ect s on the control-surface-effectiveness parameters to a qualitative indication as to t rends with Mach number.
The total known uncertainties introduced into the aerodynamic coef ficients by the factors comprising the various results were determined in the manner fully discussed in reference 5 and are of the following magni tude s :
Quantity Uncertainty
Rolling-moment coefficient ±0.0008 Yawing-moment coefficient ±O.OOOI Mach number 0.01 Reynolds number 0.03 X 106 Angle of attack ±0.05° Elevon deflection ±0.25°
RESULTS AND DISCUSSION
In evaluating the results of the present investigation, it should be noted that the tests were made a t but one Reynolds number (1. 5 million). 6 NACA RM A50r:07
Previous tests, however, made to determine the effectiveness of this elevon as a longitudinal control (reference 2), showed no appreciable change in the effectiveness and hinge-moment parameters in the Reynolds number range of 1.5 to 3.7 million.
Elevon Lateral-Control Characteristics
Results of the force tests made to determine the lateral-control characteristics of the elevon are shown in figures 3 to 8.1 Cross plots of the basic data of figures 3, 4, and 5 (figs. 9, 10, and 11) show the variation of rolling-moment coefficients with elevon deflection for constant angles of sideslip for Mach numbers of 1.2 through 1.7.
Rolling moment.- An indication of the effectiveness of the elevon as an aileron control is obtained from the rolling-moment coefficient versus elevon deflection curves of figure 9. The variation of rolling moment coefficient was generally linear with elevon deflection through 0 0 out the Mach number range for elevon deflections between +10 and -20 • 0 For negative deflections greater than 20 , however, there is an appre ciable decrease in effectiveness indicating probable increased separa tion effects with increased elevon deflection. The pitching-moment effectiveness curves (see reference 2) for this same elevon indicate a similar tendency although the lift-effectiveness curves are essentially linear indicating a redistribution of spanwise loading to give a shift in center-of-pressure position rather than a decrease in the lift effectiveness parameter.
Qualitatively, the effect of positive angles of sideslip (right wing leading) was to increase the rolling-moment effectiveness of the right wing elevon, and, conversely, negative sideslip angles decreased the rolling-moment effectiveness at a Mach number of 1 .2. (See fig. 3.) This effect of sideslip angle on rolling-moment effectiveness decreased with increasing Mach number, becoming relatively small at a Mach number of 1.7 for the range of angles of sideslip of 50 to _50. Since, in the present investigation, the significance of the effects of sideslip are restricted to a ~ualitative indication as to trends with Mach number (see section on reduction of data), and since the effects of sideslip are small over the range investigated, the CL I versus ° curves are given for various nominal angles of sideslip at 00 angle of attack only; for the higher angles of attack, data at only 00 sideslip are presentei . IThe elevon deflection angles shown in these figures (osetting) are for the unloaded condition; however, the correction due to aerodynamic loading amounted to but 10 at angles of attack of 100 and to smaller amounts at the lower angles of attack. The aileron deflection angles in the subse~uent cross plots are fully corrected for this effect of aerodynamic loading (see section on apparatus and test methods). NACA RM A50K07 7
The results from figures 9 and 11 are summarized and presented figure 12 in the form of the rolling-moment-effectiveness parameter A theoretical estimation of the variation of this rolling-moment effectiveness parameter with Mach number was made by the method of refer ence 6 and is presented with the experimental values in figure 12. As predicted by linear theory, a very gradual loss of Cl5 with increasing Mach number is evident, although only about 50 percent of the predicted effectiveness is realized throughout the speed range investigated for 00 angle of attack. This ratio of experimental to theoretical rolling moment effectiveness would be expected from the fact that but half the predicted lift effectiveness of the elevon was realized. (See refer ence 2.) Increasing angle of attack from 00 to lOo had a moderate, adverse effect on the rolling-moment effectiveness of the elevon at a Mach number of 1.2. The magnitude of this adverse effect decreased with increasing Mach number and became negligible at Mach numbers of 1.5 to 1. 7.
The significance of the results insofar as lateral control is con cerned can be assessed by using the data to determine the effectiveness of the elevons in trimming the wing-body combination to a wings-level attitude in sideslip. By the use of the theoretical lateral-stability derivatives of reference 7, it can be shown that the elevons have suffi cient rolling-moment effectiveness to hold a wings-level attitude for the flat wing of the present test to a maximum sideslip angle of between 70 and 90 at a lift coefficient of 0.25. The use of twist and camber in the wing materially improves this condition so that, for the cambered and twisted wing of reference 7, it is estimated that the elevons will hold a wings-level attitude to about 120 sideslip at a lift coefficient of 0.25.
A better quantitative indication of the adequacy of the aileron control can be obtained from the computed values of the helix angle pb/2V produced by 10 elevon deflection. The helix angles generated by the wing tips in a steady roll were calculated using the experimen tal results of figure 12 and the values of the damping coefficient Cl computed by the method of reference 8. These values of Pb!2V, thus P 5 calculated, are for the unyawed condition for 0 0 angle of attack only. Values of the damping coefficient CI p used in computing pb/2V varied from -0.252 at a Mach number of 1.2 to -0.282 at a Mach number of 1.7. The wing-tip helix angle per degree elevon deflection, thus derived, reduced with increasing Mach number, with the rate of decrease gradually lessening with increasing Mach number. (See fig. 13.)
For airplanes capable of very high speeds, maximum rolling velocity is believed to be more of a criterion of the required rolling perfor mance than pb/2V. By the use of this criterion to evaluate the rolling performance of the elevon (a rigid wing assumed) the present elevon 8 NACA RM A50K07 could be reduced in size if used as an aileron only. This supposition is substantiated by the data of figure 14 which show a comparison of the pb/2V obtained from 300 total elevon deflection with that required to attain a rolling velocity of 2000 per second with a 4o-foot wing span airplane of the present configuration flying at 60,000 feet alti tude.
The foregoing discussion of the adequacy of the elevon in pro viding lateral control must be modified in light of possible aeroelastic effects. For supersonic speeds, the loss in aileron effectiveness due to bending and torsional deformation of a swept-back wing is much greater than for the equivalent wing at subsonic speeds. The reasons for this are as follows:
Although the rolling moment due to unit aileron deflection is only about half at supersonic speed of what it is at subsonic speed (see reference 10), the loading due to aileron deflection is so distributed that the wing torsional moment (and, therefore, torsional deformation) per degree aileron deflection is about the same in both cases, provided the dynamic pressure is the same. To obtain a given C1, the aileron deflection at supersonic speeds must be about twice the deflection necessary at subsonic speed, resulting in twice the torsional defor mation. Therefore, since the rolling moment due to a given wing twist is somewhat greater at supersonic Mach numbers, aileron reversal occurs at much lower dynamic pressure at supersonic speeds than at subsonic speeds. Calculations show that, for an airplane equipped with a wing similar to that of the present investigation, designed for a maximum load fact or of 2.5 cruising at M = 1.4 at 60,000 feet altitude, the aileron effectiveness in rolling is reduced by aeroelastic effects to about 50 percent of that for a rigid wing. It appears, therefore, that in the design of ailerons for supersonic swept-wing aircraft considera tions of the effects of elastic deformation of the wing are very impor tant.
Yawing Moment.- The yawing-moment-coefficient increment resulting from positive deflection of the right wing elevon was adverse (sign of yawing-moment increment opposite to that of rolling-moment increment) at positive angles of attack. (See figs. 7 and 8.) The magnitude of the adverse yawing-moment-coefficient increment increased as the wing angle of attack increased, but was little affected by Mach number. Adverse yaw tends to retard the forward movement of the upgoing wing, resulting in a loss in pb/2V if a rudder is not used to counteract the yawing moment. However, the adverse yaw existing in the present case does not appear to be serious since estimates show that it may be overcome by use of a rudder. 2 NACA RM A50K07 9
Spoiler Lateral-Control Characteristics
Results of the investigation of 5Q-percent-semispan, outboard spoilers tested in heights of 50- and 100-percent maximum wing-section thickness (respectively, 5- and lQ-percent of wing chord perpendicular to leading edge) at 40-, 50-, and 6o-percent-chord stations showed rela tively poor rolling effectiveness. As little difference in rolling moment effectiveness was evident for the three chordwise stations inves tigated, data for but one chordwise spoiler location (5O-percent-chord station) are presented in figure 15. The maximum rolling effectiveness obtainable for the several spoilers tested was equal to that obtained from about 200 deflection of the elevon; however, a marked, adverse effect of angle of attack at angles above 40 was evident over the entire Mach number range investigated. Thus, as the angle-of-attack effect on the rolling-moment effectiveness of the elevon was only moderate at Mach numbers of 1.2 to 1.4 and insignificant at Mach numbers of 1.5 to 1.7 (fig. 12), the present elevon is markedly superior to the spoilers tested for lateral control of this wing-fuselage combination.
CONCLUSIONS
An investigation of the effectiveness of a constant-percent-chord outboard elevon and upper-surface spoilers as lateral-eontrol devices 0 for a wing of symmetrical section with the leading edge swept back 63 , in combination with a body of revolution, disclosed the following results for the Mach number range of 1.2 to 1.7 at a Reynolds number of 1.5 million:
1. Results of tests of the effectiveness of a 3O-percent-ehord, 5O-percent-semispan elevon as a lateral control indicated the following:
(a) The rolling-moment-effectiveness parameter CIa was about half that predicted by linear theory.
(b) Variation in Mach number had little effect on the rolling moment effectiveness of the elevon over the range investi gated (M=1.2 to 1.7).
(c) Increasing angle of attack from 00 to 100 had a moderate, adverse effect on the rolling-moment effectiveness of the elevon at a Mach number of 1.2, which decreased with increasing Mach number and became negligible at Mach numbers of 1.5 to 1.7. 10 NACA RM A50Y.07
(d) For a rigid wing, the present elevon is more than adequate for rolling control of this wing-fuselage combination; however, theoretical calculations showed that aeroelastic effects adversely influenced the rolling effectiveness. The elevon also appears to be satisfactory in trimming to a wings-level condition in sideslip especially for twisted and cambered wings.
(e) Deflection of the elevon produced small adverse yawing moments at positive angles of attack.
(f) The effectiveness of the elevon as a lateral control waE little affected by angles of sideslip up to ±5°.
2. Results of the investigation of outboard upper-surface spoilers, tested in heights up to the maximum wing-£ection thickness and located successively at 40-, 50-, and 6O-percent-chord stations, proved the spoilers to be inferior to the elevon for lateral control of the present configuration because of a rapid loss in effectiveness at angles of 0 attack above 4 •
Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif.
REFERENCES
1. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN 1350, 1947.
2. Olson, Robert N., and Mead, Merrill H.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 630 . Effectiveness of an Elevon as a Longitudinal Control and the Effects of Camber and Twist on the Maximum Lift-Drag Ratio at Supersonic Speeds. NACA RM A50A3la, 1950.
3. Perkins, Edward W.: Experimental Investigation of the Effects of Support Interference on the Drag of Bodies of Revolution at a Mach number of 1.5. NACA RM A8B05, 1948.
4. Frick, Charles W., and Olson, Robert N.: Flow Studies in the Asym me tri c Adjustable Nozzle of the Ames 6- by 6-Foot Supersonic Wind Tunnel. NACA RM A9E24, 1949. NACA RM A50K07 11
5. Hall, Charles F., and Heitmeyer, John C. : Aerodynamic Study of a Wing-Fuselage Combination 0 Employing a Wing Swept Back 63 . Characteristics at Supersonic Speeds of a Model with the Wing Twisted and Cambered for Uniform Load. NACA RM A9J24, 1950. 6. Frick, Charles W., Jr.: Application of the Linearized Theory of Supersonic Flow to the Estimation of Control-8urface Character istics. NACA TN 1554, 1948. 7. Lessing, Henry C.: Aerodynamic Study of a Wing- Fuselage Combina tion Employing 0 a Wing Swept-Back 63 - Effect of Sideslip on Aerodynamic Characteristics at a Mach Number of 1.4 With the Wing Twisted and Cambered. NACA RM A50F09, 1950. 8. Walker, Harold J.} and Ballantyne, Mary B.; Pressure Distribution and Damping in Steady Roll at Supersonic Mach Numbers of Flat Swept-Back Wings With Subsonic Edges. NACA TN 2047, 1950. 9. Frick, Charles W., and Chubb, Robert S. : The Longitudinal Stability of Elastic Swept Wings at Supersonic Speeds. NACA TN 1811, 1949. 10. Jones, J. Lloyd, and Demele, Fred A. : Aerodynamic Study of a Wing Fuselage Combination 0 Employing a Wing Swept Back 63 .- Effects at Subsonic Speeds of a Constant-Ghord Elevon on a Wing Cambered and Twisted for a Uniform Load at a Lift Coefficient of 0 . 25. NACA RM A9I27, 1949.
Aspect ratio 3.50 ~ Hinge line 10-percent chord Taper ratio 0.25 ~ ~ 'c!j Airfoil section perpendicular ~ -..:] to leading edge: NACA 0010 I 17.25 L ~~
- .<,''-''> .... ,," ,,// .... ' ---b-- ...... rmox =i 38 I"~', ------~ .... J- , ------~J
I 7.00 L _I. .57 • /6.50 ~...... _- " ,-L I /2 35.00 /3.14
Dimensions in inches
59. 50 ~4.00~j I.- ~ Figure I. - Sketch of model showing principal dimensions and location of elevon, I-' W
I
~
§1
:x> \J1o i:3 -.J
Figure 2.- Model installed in the wind tunnel showing elevon with negative deflection.
f-' \J1
3 NAeA RM A50K07 17
.02
.01 b. o b. io- ... - .01
ZJ- .02
.A-
t"-
()...
8 Setting (b) M =1.3. o 9.80 o -/. I 0 <> -9.70 l::,. -20.50 Ll -30.1 0
.01
o G-
-.01 -6 o 2 4 6 Angle of sideslip, /1', deg """~--:NA""'CA=-""" (c) M~1.4 . Figure 3. - Effect of elevon deflection on the variation of rolling-moment coefficient with angle of SIdeslip at various Mach numbers. 0=0 0 18 NACA RM A50K07
.01
o I h '"
-.01 (d) M=/.53.
"
v
'"U"
(e) M=/.6. 8 Setting o 98° o _1./0 <> -9.r 6 -20.5° .011 -30.1 °
.01 1---.
o
L_
-.01 -6 -4 -2 o 2 4 6
Angle of sideslip I #', deg ~~-NA-----CA~ (f) M=/. 7.
Figure 3 - Concluded [- NAeA RM A50KD7 19
.02 :--. ~ r-.. ~ r- ...... :>-...... I---.:t-- .0 1 ""'- ""'- "-...... v- I--- r-- ~ ...... -r--- "6... r--, r--...... r--... r--- r- ~ J...... ~ - 0 ...... ""b-. r--. r--- :"--.. r-... ~ . r--..... ""0..... -.0 1 ~
-.02 ( a) M =/2.
~- .02 ,-' t:: ~ . ~ ~ r-- r-. r- ;.::: .01 - ~ ..... r---,~ 'b ~ r--~ I:) r-... r-- ~ ~ ~ -0-. r-- --r-. ~ r...... r-. ~ r-- -0.. ~ '- 0 ...... ~ r--. r--. ""'-u;.. e:: r-I -.... t--. I:) ...... - e:: - .0 1 ...... I ...... tl\ .!:: ...... ~ -:02 (b) M =/ 3 .
.02
~ .0 1 ~ ~ ~ t-- ....::: Cl-- r-o-- - F==:: t:-N... r--. --- r- ~ r--~ r-::::I-- 0 r- ""----- S::! -<>.... - "-tr- r--...... - r-- r- - r- -: 0 1 f--_ 8 Se tfi% --...,. 0 9.8" f), - .5° _0-1./° Lr30l o f-- -- <>, 9.7° 1 1 -.02 1 -6 -2 o 2 4 6 Angle of si deslip, /f', deg (c) M =/ 4 . "'""~---:N-'-A=(A--?
Figure 4 . - Effect of elevon deflection on the variation of rolling moment coefficient wllh angle of sIde slip at various Mach numbers. a =5 ° 20 NAeA RM A50K07
.02
t-- .01 ~ r&- t-r---~ r:::::: ~ ~ t-- r--- --t--r-r--I'-n-. r--.: o r r-a.. r0- t-- ~ I- r--~ t-- 1'--) t-- ,- r-a.. I--- :-- --- r---0... -.0/ - -....
-.02 (d) M=/.53.
~- .02
~ r--::: F:t ~- t--. ~ to-r---~. v r-- ~ t--t---t-- ~ t--!-n- r---t-- ~ t-- ...... o r-- t---0.. r--~ ~ - r-e- --- g -- -- i"'-t-- - E: -.01 ...... I l:), -- --- .C:: -::::-..... ~ -.02
.02 8 Settinlt, 0 09.8 0 f:, - 0.5 0-/./ 0 Ll -30.1 0 .01 0-9.r - I-I- ">- -.... -- J..- r-- ~ --.; o ,.-. v ~ r-- - -r-- -0..r-- -. r--t- "-Q... -.01 r--t--- -6 -2 o 2 4 6
Angle of sideslip I ,e', deg (f) M=/. 7. "'"~~NAC-:C"""A~
FiQure 4. - Cone luded. NACA RM A50K07 21
.02 t.. .;-r---~ Ir .01 [. ~ ...... v - ...... ~ ::::-- ~ I- ...... - - r-;;::::: t=:;:1.r-- '- ~ o ~ - ...... ~ r--..... "b. --...... -.0/ '-...c
~- .02 ...... ' c::: . ~ ~ .~ ¢.. ::::::: ~ :::: .01 Q.) ~ ~ ~ ~ :::::::: r==-- """'- (.) ...... ~ ...... ~ -- ~-.- ""'0 ~ ...... ~ -- r---.. '""0...... I'------0 ~~ .... I'--c
(b) M=/.3.
Angle of sideslip, #', deg (c) M =I. 4. ,,<="~--:ONA"""'CA-:-
Figure 5 . - Effect of elevon deflection on the variation of rolling moment coefficient with angle of sideslip at various Mach numbers. a = 10° 22 NACA RM A50K07
-.0 1 1-----+----t--+---+-+---+-+--+---''''''''''"::::-+--I----1
(d) M=1.53. ~- ...... ' t:: . ~ ..;::(,,) .01 k ...... J". -..::::::::::
.01
8 Setting 0 0 09.8 /::;. -20.5 -.01 o -1/ 0 LJ -30. 10 -+--f---+--+-----+-''''''''-Io:::::--+----i 0-9.r -6 -4 -2 a 2 4 6 Angle of sideslip, ,11', deg ~ (f) M=17
Figure 5. - Concluded
. . ------~ NAeA RM A50K07 23
.008
tli... ~ .004 ~ ~ j-...... ~ t---.,... - h ...... ~ ...... --~ " ~ Il1o.. 0.. , o ...... ~ ~ ...... ~ i'-. ~ r---...... ~ to-... ~ In_ - II:: :004 ~ ~ , ~ ~ ...... ~ c:::: r . ~ -.008 ~ (a) M=/'2. -....:~
A"gure 6 . - Effect of e/evon deflection on the variation ofyawing moment coefficient with angle of sIdeslip at various Mach numbers. a = O~ NACA RM A50K07
.008 i'---- ...... ~ '"...... - I'...... 004 ...... r-c...... ~ ~ ~ "- ...... ; ~ ~ ~ '" I'-...... o i'O.: , r--...... ~l...... '"""0... ~ ~ ...... ~...... ~ '" ~ ... -.004 ~ ~ ~c:: I" .. ~ $ .-~...... -.008 (,.,) ,~ (c) M=/'4.
-6 -4 -2 o 2 4 6 Angle of sideslip, fi', deg (d) M=I.S3.
Figure 6. - Continued. 4 NACA RM A50K07 25
~
~ ~ .008 ...... ll...... r-...... r-...... lI.. -a...... -...... ~ ~ r-...... ~ .004 ...... ~ r-...... ~ ~ ...... --...... :: ...... t::::: ~ r...... ~ ~ I':::::" r (e) M=/. 6.
-~ ~£:: .012 I ~ , ~ . ~ ...... ~ ~ ~ ~ ~ .008 ...... - ... ~ ~ ~ ...... ~ N ~ .004 ~ ~ r..... "" ~ ...... I'--...)L ~ ~ i'G...... ~ ...... o ~ ~ ~ ~ Setting ~ ...... r-- 8 . r,J 0980 D. -20.50 0 0-1./0 LI-301 ' 1- -.004 o O-r.7 , I I " -6 -4 -2 o 2 4 6 Angle 01 sideslip, #', deg
(I) M:~l.7. ~
Figure 6. - Concluded. 26 NA9A RM A50K07
.004 b-f-.... ~ -0.... I--. o ~ ~ ...... ~ ::---- ~ ~-...... ~ ~ ~ -.004 ~ ~ =::::::::-.t -----0:.:: ~ ~I 'i>
-.004 8 Setting 0980 f':l -20.5 0 o -1./ 0 L1 -30./ 0 -+------If------+--f------+------I~....,-__I 0-9.70 -.OOBL-----"----+--'----L--L----L-----1.--'---L-L----L-----1 -6 -4 o 2 4 6 ~ Angle of sideslip, ,0', deg (b) M=1.3.
Figure 7 - Effect of elevon deflection on the variatioh of yawing moment coefficient wtlh angle of sideslip af various Mach numbers. a = So. NACA RM A50K07 27
, ~~004~~~--4--+--+-~--~~~~~~~ ~
...... ~~008~~~--~~--~-L--~~~--~~~ .(.) ~ (c) M =/,4. 'to..:
...... ~ ~ .008 ~ ~ .~ ~ .004 ~ ~ ~ ~ ~ ...... ~ ~ o ~~ ~ I - r...... : ~ ~ ~ ,.... ~ ~: -.004 r- 8 Setti~ ~ ...... 098° D. - 0.5° ~ r- 0 _/./0 Ll -30./° ~ 0-9.7° -.008 I I I I -6 -4 -2 o 2 4 6
Angle of sideslip I ,8', deg (d) M=/.53
Figure ? - Continued. 28 NACA RM A50K07
~c::
~"'-004 ~---1...... _L...----1...... _L...----1...... _&'---L..._.1----L..._.1----L...----'
o 8 Setting o 98° ~ - 20.5° 0-1./ ° Ll -30. 1° -+-+--+-+--+---"~,,",*"-----l 0-9.7° -.004 L...----'------I_~___6__-'--__&__L....___'____'__...J_____'_____l -6 o 2 4 6
Angle of sideslip, ,e'l deg
(f) M=/.7
Figure 7 - Concluded NACA RM A50K07 29
. .004 ~ --0...,. o ~ ~ ...... ~ ""-.... """a.... ~ , r-~A. v...... ~ ~- ~ ...... - - , I--.. :004 ., ... ~ ~ ~ (.:)c:: ~ ~ ~ ...... -:008 ~ ~ (a) M-=12. .~ ,'-~
0
-«J4 8 Setting 09.8° t::,. -20.5° 0-1./° L1 -30.1° 0-9.7° -.008 -5 -4 -2 0 2 4 6 Angle of sideslip, p', deg ~ (b)M=13.
Aqure 8. -Effect of elevon deflection on the variation of yowing moment coefficient with angle of sideslip ot various Moch numbers. a=/O~
J 30 NACA RM A!)OK07
.004 r'. t--... ..,...., o ~ ", ~ -n... '" ...... ~ r-...... r...... ~ ...... ~ . ~ -~ ~ .~ ~ J.. ~ ~ -:004 ~ ~ .... '" ~ ~ "" "'G
(c)M=1.4 .
...... t:::
-:008~~~-~~-~-L-~~~-~~~ -6 -4 -2 o 2 4 6 Angle of sideslip I ,8'1 deg (d) M=/.53
Figure 8.-Continued NACA RM A50K07 31
I - .008 ~ ...... 004 ~ K r-...... ~ ~ ~ F:::::!a. ~ ~ ~ ""'""'" ~ ...... ~ ~ ~~ o -.....;;::: ...... ~ ~ ~v...... , ~ ~ 1- ~~ -.004 "'IIi
(e) M=16
-~ I I I I ~ 8 Setting _ .008 0 ~ ~ 098 .6. -20.5 ° I ~ ~ L.1 -30.1 0 _ tr, 0-1." . ~ ~ 0-9.7° ~ ~ ~ ...... ~ .004 «: ~ ~ ~ ~ "-~ o ~ .... ~ ~ "~ ~ ~ -.004 ~ ~ -6 - 4 -2 o 2 4 6 Angle of sideslip, ,8', deg
(f)M=f1
Figure 8. - Concluded LA! f\)
~-.Oll 1 ...... jJ .0/1 I ...... , , .0/1 I ...... !'
...... ~
~
~ 0 lac:: I >k:f:t k>l<>1 z' I> cl:: o 1 1 > 4:::P ol ?>otcit:."I <' P I _" K::> Ie "'I < o l 1>.c t cd > '" > k "'4< :§WtffFtj > ~
Elevon deflection, 8, deg Elevon deflection, 87 deg Elevon deflection7 8, deg (a}M-=12 (b) M =13 . (c)M =14 .
Figure 9. - Variation of rolling -moment coefficient wdh elevon deflection at various Mach numbers. a= 0 ~
~
~ :r> Vlo ~ -...J 5 NACA EM A50K07 33 I Q
~ ~, G<;:) r:::;' , ~ "'- 't)-..: C) ...... Q) " I ~~ ~~ '- § C) :::.. C\a I ~ ~
...... g C) C) C) C) C) C) (3' " ~ 01 'ti ~ ~ C)~ ~ ~ §' ~ § :.;::...... : (..) ~ C) " ~ ... ~ ~-.. o'i ~ ~ Q) r:::; , ~ ~ ~, :::.. ,!::h ~ ~ ~
C) ...... 1'1) C) C) C) C) C) C) C)' " C)...... !::h ~ , C) ~ r:::;' 1'1) ~\{) :0:::: .....: ~ ~, Q) " it::: ~ ~ ~ '-
C) ...... C) C) C) C) (3~ " ,0 'IUBl:J/jjBO:J I UCiWOW-OIJ///OC/ I..AJ -F
.02
I'---. .01 r---..... r----... """-- (..:)'" - i'---. r--r--... r-- ...... -- --- ....1'---...... i'---. I'--...... I"-...... I"-...... ~ 0 ...... :--...... ~ l'--... ~
Q.l I '"(;) -.01 '" ~- '----- ~ (.) (a)M=/.2 (b) M::-/.3. (c) M : 1.4 ......
~
~ E: I 1::1\ ,~ .01 (;) Q:: ---- o -r- r-...
I'--...... Nttfttl ffftf4tl ~ -.01 -30 -20 -10 o 10 -30 -20 -10 o 10 -30 -20 -10 o 10 E1evon deflection, 8 , deg E1evon deflection, 8, deg E1evon deflection, 8, deg (d)M= I. 53 . (e)M=1.6 . (f)M-=-!.? ~ CJ ;t>
~ ~ Figure 10.- Variation of rolling-moment coefficient with e/evon deflection 01 various Mach numbers. a=5° , jJ:'o°. ;:t> \Jlo ~ --J ~ !xl :s:: ~ .02 .02 .02 'c!:l ~ -..l r-- .01 .01 .01 - ...... - I'-.. ~ ...... - ...... --- r-...... "- ~- 0 o o - ...... "- i'--...... ~ .Cb
.~ -:01 -.01 -.01 ,~ Cb I:) (a) M=/'2 . (b) M'/ 3 . (c) M=14. ~
.....
~ ~ I:)
;:1r------;--r--, Hf~§J 1f O~IIIi#HJO~f1 ffHtll -30 -20 -10 0 10 -30 -20 -10 o 10 -30 -20 -10 o 10 Elevon deflection, 8, deg Elevon deflection, 8, deg Elevon deflection, 8, deg (d)M :: /.53 . (e)M = /.6 . (f)M =/' T.
0 Figure II . - Variation of rolling-moment coefficient with elevon deflection at various Mach numbers.~ IOD ,IJ~ a = , 0 .
w \Jl lA.l 0"
.003 Linear Theory ~ -- Experiment ~
~ ~ .002 I- _ r-- r- - o -- -- - ~- ~- C) -. t--- ~ II ------~ .... , 001 a=O° (.)..... ~ . '- /0° !-'
o ~ 1,/ 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 Mach number, M
Figure 12. - Variation with Mach number of the rol/ing-momenf- effectiveness parameter, C, . Data for differential deflection of two elevons. /1'= O~ ~ 6 ~
~ ~ \.Jlo -:]~ MeA RM A50K07 37
,~ ~ ~
~ Cb o. ~C) § II ~~ t:~ ..c:::U . ~ CI) ~~ ~ ..c:::~ ~Cb ~ <::) Cb ~ ...... ' ~ ~ t:' ",'" ~ <::) Q) ~ ..::::.~ ~ § q,=t::: ~ ..c::: U Cb .~ j ~ ;;:: "':'-~ ~ ~ .~~ / ~ ,..... Cb t: ~ ,..c::: ,Cb ,Cb <::)::::..... / t:'=» .~ / .~",=, <::) ~
~
't (\j ~ - ~ ~ \.) - .~ ~. ~ . a:: SUD! pDJ ' e/t1~/qd Lv ()) .10
.08
CI) c::: ....t) t) ~ ""- .06 0 0 +100) ~ 6TOTAL =30 (-20 8 - ---.r-- ... ~ r:--:: ..... ~ ...... .02 . o ~ I 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 Mach number, M ~ ~ Figure 14. - Comparison of the wing-tip helix angle attainable wltll ~ 0 30 total deflection 'lie elevon necessary to produce a ~ of wlt"'"at \Jl wing- lip velocily·of 200 0 per second for a 40-fool wing span o ~ airplane flying at 60,000 feet. a = 00. -...l NACA EM A50K07 39 .01 ------0 -~------~- ~- --~ -- ~...... (o)M =/. 2· ...... c:: .~ .~ ::::Cb 8 /JI ...... - ~ ~ .---,...... ~ 0 - .--... -.-_ ~- I------.. -- ~ .~ § c::s '-. (b)M=I.53. ~ ...... ~ .01 ~ a.:/: ~ 4 ------8 0 1--- P-- I-- - ~- . o -" ------o 20 40 60 80 /00 Spoiler projection, perc8l7t maximum wing-section thickness (c) M =1.7 Rgure 15. - Effect of upper surface spoIlers on rolling -moment coefficient for various angles of of/ock ond Mach number. Spoiler on one wing only. NACA-Langley